orbitalElements
Orbital elements of satellites in scenario
Description
Examples
Retrieve Orbital Elements of Satellite
Create a satellite scenario object.
sc = satelliteScenario;
Add a satellite to the scenario.
tleFile = "eccentricOrbitSatellite.tle";
sat1 = satellite(sc,tleFile);
Retrieve the orbital elements of sat1
.
elements1 = orbitalElements(sat1)
elements1 = struct with fields:
MeanMotion: 1.4544e-04
Eccentricity: 0.7415
Inclination: 60.0000
RightAscensionOfAscendingNode: 30.0000
ArgumentOfPeriapsis: 280
MeanAnomaly: 289.4697
Period: 43200
Epoch: 05-May-2020 13:51:55
BStar: 0
Add a satellite from Keplerian elements to the scenario.
semiMajorAxis = 6878137; % meters eccentricity = 0; inclination = 20; % degrees rightAscensionOfAscendingNode = 0; % degrees argumentOfPeriapsis = 0; % degrees trueAnomaly = 0; % degrees sat2 = satellite(sc,semiMajorAxis,eccentricity, ... inclination,rightAscensionOfAscendingNode, ... argumentOfPeriapsis,trueAnomaly, ... "OrbitPropagator","two-body-keplerian", ... "Name","Sat2");
Retrieve the orbital elements of sat2
.
elements2 = orbitalElements(sat2)
elements2 = struct with fields:
SemiMajorAxis: 6878137
Eccentricity: 0
Inclination: 20
RightAscensionOfAscendingNode: 0
ArgumentOfPeriapsis: 0
TrueAnomaly: 0
Period: 5.6770e+03
Input Arguments
sat
— Satellite
row vector of Satellite
objects
Satellite, specified as a row vector of Satellite
objects.
Output Arguments
elements
— Orbital elements
structure
Orbital elements of the input sat
, returned as a structure. The
fields of the structure depend on the orbit propagator you specify using the OrbitPropagator property of the satelliteScenario
object.
Two-Body Keplerian
The two-body-keplerian
orbit propagator returns these fields.
SemiMajorAxis
, in metersEccentricity
Inclination
, in degreesRightAscensionOfAscendingNode
, in degreesArgumentOfPeriapsis
, in degreesTrueAnomaly
, in degreesPeriod
, in seconds
SGP4 and SDP4
The
sgp4
and sdp4
orbit propagators returns
these fields.
MeanMotion
, in degrees/secondEccentricity
Inclination
, in degreesRightAscensionOfAscendingNode
, in degreesArgumentOfPeriapsis
, in degreesMeanAnomaly
, in degreesEpoch
BStar
, in 1/EarthRadiusPeriod
, in seconds
The orbital elements represent the mean values at
Epoch
.
Ephemeris
The
ephemeris
propagator returns these fields.
EphemerisStartTime
EphemerisStopTime
PositionTimeTable
VelocityTimeTable
GPS
The
gps
propagator returns these fields.
PRN
SatelliteHealth
GPSWeekNumber
GPSTimeOfApplicability
, in secondsSemiMajorAxis
, in metersEccentricity
Inclination
, in degreesGeographicLongitudeOfOrbitalPlane
, in degreesRateOfRightAscension
, in degrees/secondArgumentOfPerigee
, in degreesMeanAnomaly
, in degreesPeriod
, in seconds
The orbital elements are derived from the SEM almanac file and defined in the Earth-Centered-Earth-Fixed (ECEF) frame.
Version History
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